1. Field of the Invention
The present invention relates to a shroud segment for a turbine stage of a gas turbine engine
With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate-pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
FIG. 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.
Internal convection and external films are the prime methods of cooling the gas path components—airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the working gas annulus endwalls, which include NGV platforms 33, blade platforms 34 and shroud segments 35 (also known as shroud liners). However, the flow of air that is used to cool these endwalls can be highly detrimental to the turbine efficiency. This is due to the high mixing losses attributed to these cooling flows when they are returned to the mainstream working gas path flow.
One option is to cool the platforms and shroud segments by an impingement flow of cooling air on the back surface of the gas washed wall of the component. For example, a perforated plate spaced from the gas washed wall and supported by pedestals can form impinging jets, and the spent coolant can then flow back into the working gas path at the rear edges of the component. Unfortunately, limited numbers of impingement jets can produce non uniform heat transfer distributions, and the cross flow from spent coolant can reduce the effectiveness of the impingement jets at the more downstream locations of the component. In addition, the need to keep the coolant pressure at a level above that in the working gas path reduces the allowable pressure drop across the impingement jets, and hence the associated heat transfer levels.
Thus, in the case of shroud segments, there has been a move towards the use of abradable coatings that provide a thermal insulating barrier on the gas washed surface of the segment. The corresponding blade tips may have abrasive coatings attached in order to facilitate the cutting of a track into the abradable coating. These coatings have proved effective at reducing the heat flux into the segments. However, their low thermal conductivities introduce a high thermal gradient across the thickness of the coating. Consequently the gas washed surface becomes very hot, and if not protected can increase to a temperature exceeding the sintering temperature limit of the coating material. Similarly, the bond coat that typically attaches the abradable coating to the segment also needs to be kept below a certain temperature to prevent the interface between the coating and the bond coat alloy from oxidizing and prematurely shedding the coating.
2. Description of the Related Art
In an attempt to reduce the temperature of the gas washed surface of these coatings, a film of cooled air can be provided between the hot working gas and the coating. This can be achieved by the introduction of effusion cooling onto the surface of the segment. FIG. 3 shows an isometric view of a typical shroud segment with an abradable coating and effusion cooling. The segment has a cast alloy body 40, mounting legs 41 for mounting to the turbine support casing, upstream 42 and downstream 43 edges, a feather or strip seal leakage control groove 44, an abradable surface coating 45, and a plurality of effusion cooling holes 46 which deliver streams of cooling air onto the gas washed surface of the surface coating. These streams of cooled air form a protective film on the gas washed surface. However, the film mixes with the hot gas that is adjacent to the surface and progressively heats up as it flows over the surface. This degradation of the film is normally expressed as a “film effectiveness”, which typically deteriorates with distance from the holes.
To try to achieve a high level of film effectiveness, it is conventional to introduce the cooling air onto the surface with a low momentum in order to match the momentum of the gas in contact with the wall. If the blowing rate of the film is too high then the film will blow off the surface and mixing between the gas and coolant will be encouraged. The holes can be formed as fan shaped openings in order to diffuse the flow as it exits the hole. The angle of the holes' feed passages is also an important parameter. In general, a shallower angle of the line of the feed passage relative to the gas washed surface (i.e. a smaller radial angle relative to the engine axis) helps to prevent the film from becoming detached from the surface.
Although the gas washed surface of a shroud segment has a close clearance to the tips of the turbine blades, a leakage flow of working gas nonetheless passes through the gap between the blade tips and the segment. This leakage flow is detrimental to engine efficiency, and also sets up a pressure gradient that is a destructive influence on the cooling film. The present invention is at least partly based on the realisation that shroud segment cooling air can be used to reduce the amount of such leakage flow.